Rocket propulsion systems at present employ either a pressure fed System, hereinafter referred to as a propellant tank pressurization system, or a turbopump system to transfer propellant to the combustion chamber. With the turbopump system a turbine driven pump is used to supply the propellants to the rocket's engine thrust chamber. This turbine is driven by either a gas generator or "expander cycle". The alternative propellant tank pressurization system pressurizes the main storage tanks. Under the force of that pressure, the storage tank discharges stored propellants into the engine thrust chamber.
Each of the foregoing fluid feed systems are capable of operating in a micro gravity or adverse acceleration environment, which is commonly found in upper stage rockets and space vehicle operations. Those feed systems, therefore, require sophisticated propellant acquisition devices unique to those environments. Although useful, those prior feed systems have a number of known disadvantages.
The disadvantage of a turbo pump system is its high cost, limited life, and, in certain situations, its reliability. Turbo pump systems are expensive to develop, procure and maintain. In addition, the extraction of hot combustion gases, the expansion gases necessary to drive the turbopump turbines, increases system complexity over the alternative propellant tank pressurization system and also reduces the turbines reliability.
Moreover, if a turbo pump is being used for a hybrid system, one that uses a solid fuel and a liquid oxidizer, several additional disadvantages are realized. First, no liquid or gaseous fuels are available to supply a gas generator driven turbine; only the oxidizer is available. The combustion gases in the main thrust chamber exceed the maximum operating temperature of existing turbine materials. That temperature effectively eliminates those gases as a turbine drive fluid, unless a complex method of moderating such high combustion temperatures is later developed. Further, a reliable, simple and effective method to start this type of system, ie. "boot strapping", is unavailable.
If instead, a gas generator cycle is selected, then additional fuel, such as RP-1, liquid hydrogen (H.sub.2) and the like, must be stored on board the vehicle to generate the desired combustion gas conditions for driving the turbo pump turbine. So doing increases the cost and complexity of the fuel delivery system.
A regeneratively cooled nozzle could also be used to drive a turbo pump turbine. However, since hybirds are selected based on the operational simplicity characteristic of conventional solid rocket motors, the addition of a regeneratively cooled nozzle necessarily lessens the attractiveness of the hybird system. Further, if, despite the greater weight and complexity, a regeneratively cooled nozzle is selected nonetheless to drive the turbine, the expansion gases must be transferred from the nozzle section to the turbo pump turbine, which is located above the combustion chamber, then returned and discharged to either the combustion chamber, nozzle or overboard vent. Those requirements further increase cost and complexity.
Turbo pumps are not available for work in applications having extremely low flow rates, such as existing in presently proposed electric propulsion system designs.
Turbo pump systems face severe operational issues when used for long term applications involving high temperatures as result from LH.sub.2 /LO.sub.2, or involving corrosive fluids, such as nitrogen tetraoxide or nitric acid; in particular, turbo pump systems must be essentially completely disassembled and refurbished for each flight with reusable launch vehicles, such as the shuttle orbiter.
The prior propellant tank pressurization system has singular drawbacks. Although simple in structure, propellant tank pressurization systems impose a significant weight penalty associated with pressurizing the main propellant tanks. In order to confine the high pressure levels, the tank must be strong. That strength is equated with greater thickness of the pressure vessel material, such as steel, aluminum and the like metals, and, accordingly great weight. As higher pressurization levels are specified, the weight of the tank increases. Not only is weight of the tank a penalty, but, as larger pressures are specified, a practical limit is reached to the tank size and weight for the rocket application. To minimize the tank weight, as a practical matter, the combustion pressures must be lowered in order to reduce tank supply pressures. This results in reduced performance, as compared to a typical turbo pump feed system. As a consequence, low thrust chamber pressures, and, hence, reduced performance, must necessarily be accepted in order to maintain an acceptable tank weight and corresponding booster mass ratio, the effective propellant mass divided by the initial vehicle mass.
To acquire propellants in a micro gravity or adverse acceleration environment, propellant acquistion devices are required. Those devices are designed to use the surface tension characteristics of the propellants to maintain a continuous, uninterrupted supply of propellant flow. Such acquisition devices for propellant tanks are necessarily large in size and are therefore much more susceptible to the hydrostatic and dynamic forces that result from adverse acceleration fields. Those forces can cause the propellant acquisition device to "breakdown" or, as variously termed, "breakthrough", resulting in an undesirable two-phase flow condition in which gas is ingested into the turbopump or engine, causing pump damage from cavitation and combustion problems. That situation is aggravated when large propellant tanks contain low quantities of propellant, as occurs when a major portion of the stored propellant is depleted. With the lesser volume of propellant, the surface area of the screen or channel in continuous contact with the suspended liquid, such as occurs during engine re-start, is reduced and thus the pressure drop of the liquid flowing through the screen is increased, and can lead to "breakdown", namely gas ingestion.
Other techniques for acquiring propellant for engine re-start operations require vehicle rotation or linear acceleration to settle the propellant and keep it at the bottom of the tank. Those approaches are also known to have adverse performance impact on space vehicle design in that they require additional propellant to accelerate the vehicle with small thrusters before the main engines can be started.
An object of the invention therefore is to provide a novel and versatile fuel delivery system.
An additional object of the invention is to provide a fuel delivery system for rockets that avoids the disadvantage of a turbo pump system or the weight penalty associated with main propellant tank pressurization in the propellant tank pressurization system.
A further object of the invention is to provide a fluid transport system that is of lesser weight and improved performance than the prior single tank pressure fed system.
A still further object of the present invention is to provide a fluid transport system that is of greater maintainability and reliability than a turbo pump system.
A still additional object of the invention is to reduce the problem of propellant acquisition in a micro gravity or adverse acceleration environment.